The present invention relates to combustion chambers for aircraft type turbojet engines, specifically such combustion chambers having a generally annular configuration with a double walled converging zone.
It is well known that the power output of an aircraft type turbojet engine is directly related to the temperature of the combustion gases at the inlet of the turbine section of the engine. In order to increase the power, today's jet engines have such operating temperatures on the order of 1,800.degree. K. However, the specific fuel consumption of the jet engines is very high at these operating temperatures. In order to maintain acceptable fuel consumption, present day engines have been designed with increasing compression ratios.
While the increased operating temperatures and the increased compression ratios have provided jet engines with improved performance and specific fuel consumption figures, the structure of the combustion chambers has had to be modified to obtain the necessary strength and the requisite service life. The combustion chambers are typically designed to incorporate double thickness walls on both the inner and outer annular boundary walls so as to provide the necessary cooling to the inner walls, which are exposed to the combustion gases. Failure to provide such cooling will typically reduce the service life of the combustion chamber or result in structurally weakened areas. The provision of the double walls throughout the length of the combustion chamber has increased both the complexity and the inherent costs of such turbojet engines.
French patent Nos. 2,567,250 and 2,579,724 illustrate the known type of double-walled combustion chamber wherein the hot wall portions are retained in the adjacent cold wall portions by the interengagement of a lip formed on the hot walls with a groove formed in the cold walls.
Typical of the prior devices, these patents also show the attachment of the downstream end of the hot and cold wall to the turbine inlet area by means of one or more bolts passing through corresponding flanges. This rigid attachment does not provide for any relative radial expansion between the combustion chamber and the turbine inlet.